1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to a platform and blade assembly for use in a turbine of a gas turbine engine.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
Rotor blades in an axial flow compressor or turbine used in a gas turbine engine have a rotor disk with a plurality of dove-tail or fir-tree slots formed in the disk in which a blade root having a similar cross section shape is placed in order to secure the blade to the rotor disk and hold the blade against the high centrifugal forces that develop during operation of the engine. The turbine blades typically include platforms that extend between adjacent blades and form an inner shroud for the gas flow through the blades. Stresses induced by the high rotor speeds concentrate at the fir tree slots and can be minimized by minimizing the mass of the blade.
Nickel base super-alloys are widely used in applications where high stresses must be endured at elevated temperatures. One such application is the field of gas turbine engines where nickel base super-alloys are widely used especially for blades and vanes. Demands for improved efficiency and performance have resulted in the operation of turbine engines at increasingly elevated temperatures placing extreme demands on the superalloy articles used therein.
One approach to improve the temperature capabilities of nickel based super-alloys is to fabricate the blades in the form of single crystals. Conventionally prepared metallic materials include a plurality of grains which are separated by grain boundaries which are weak at elevated temperatures, much weaker than the material within the grains. Through specific casting techniques, nickel based super-alloys can be produced in single crystal form which have no internal grain boundaries. U.S. Pat. No. 4,719,080 issued to Duhl et al on Jan. 12, 1988 and entitled ADVANCED HIGH STRENGTH SINGLE CRYSTAL SUPERALLOY COMPOSITIONS shows a prior art single crystal turbine blade, the entire disclosure of which is incorporated herein by reference. A single crystal blade will have higher strength in the radial direction of the blade which will result in better creep strength and therefore longer blade life.
Recent casting technologies have made the casting process for a single crystal blade at about the cost of casting a non-single crystal blade. However, casting process for single crystal blades produces a larger number of defective casts than does the non-single crystal casting process. This results in the casting process for the single crystal blades to be much higher. One major reason why this is so is that the single crystal blades are cast with the blade platforms formed with the airfoil portion. The platforms extend from the airfoil portion at substantially 90 degree angles from the blade spanwise direction. Since the single crystal orientation is along the spanwise direction of the blade (to provide for the higher blade strength and creep resistance), extending the single crystal growth of the blade airfoil out along the platform results in a lot of defects in the casting process. It would be beneficial to therefore from a single crystal blade with the platform formed separately in order to decrease the number of defective single crystal blades.
In some prior art turbine rotor disks used in gas turbine engines, the turbine blades have been formed from ceramic composites in order to allow for higher gas flow temperatures in the turbine section. The ceramic blades were formed with fir tree shaped roots for insertion in the fir tree slots of the metallic rotor disk. However, this manner of securing the blade to the rotor requires the blade root to be capable of withstanding high tensile forces. Ceramic materials are capable of withstanding high compressive forces, but not high tensile forces.
The prior art U.S. Pat. No. 5,030,063 issued to Berger on Jul. 9, 1991 and entitled TURBOMACHINE ROTOR discloses a rotor for an axial flow compressor or turbine in a gas turbine engine in which the rotor disk includes a plurality of fir tree shaped slots in which a turbine blade is secured within, and a ring that has airfoil shaped slots in which the blades extend through so that the ring forms a cylindrical platform for the gas flow through the blades in the assembled rotor disk. The ring an annular short flange and an annular long flange integral with the ring and on opposite sides of the cylindrical platform. The Berger invention separates the platforms from the blades so that the radial forces acting on the platform are transferred to the rotor disk instead of through the blades. However, in the assembly is used in the turbine section of a gas turbine engine, the extreme high temperatures would produce high thermal stresses on the annular flanges that would shorten the life of the ring. The lower edge of the annular long flange would be exposed to about 700 degrees C. while the upper edge would be exposed to about 1200 degrees C., resulting in a temperature gradient in this part of about 500 degrees C. which would cause very high thermal stresses in the part.
It is therefore an object of the present invention to provide for a turbine rotor disk with a single crystal blade with a platform formed as a separate attachment to the blade in which the thermal stresses would be acceptable for low cycle fatigue (LCF) and longer life.
It is another object of the present invention to provide for a turbine rotor disk with blades made from a single crystal superalloy with a lower number of defective blades made in the casting process.
It is another object of the present invention to provide for a turbine rotor disk in which the rotor blades are made from a ceramic material and attached to a rotor disk made from a metallic material, in which the ceramic blade is secured to the rotor disk and blade platforms through compression forces with very little tensile forces.
It is another object of the present invention to provide for a turbine rotor disk with blades made from a single crystal superalloy with a platform separate from the blade and secured to the blade through a shear pin that also provides for a seal between the airfoil and the platform against the hot gas flow.